1. Field of the Invention
This invention relates generally to a load carrying tubular member and, more particularly to a tubular member that has been wound with specific cross-section fibers in a controlled orientation to optimally carry the load applied to the member, where a number of load carrying members can be assembled to cooperate in forming a body of revolution.
2. Description of the Prior Art
Over the past two decades, the use of fiber composite materials in aircraft structures has gained popularity. As a result, modem airframes incorporate structural components made of composite materials to form aircraft wing structures, rotor blades, fuselage segments and the like as substantial weight savings can be achieved due to the superior strength-to-weight ratio of fiber composite materials as compared with the conventional materials of aircraft construction such as metal alloys. By replacing structural components previously formed of metal alloys with similar versions of the same component formed of composite material, a respective weight savings in the order of 25 to 30 percent is generally considered to be achievable.
In general, composites include a reinforcing material suspended in a “matrix” material that stabilizes the reinforcing material and bonds it to adjacent reinforcing materials.
Composite parts are usually molded, and may be cured at room conditions or at elevated temperature and pressurized for greater strength and quality.
Most of the composites used in aircraft structures comprise of filament reinforcing material embedded in a polymer matrix. A primary advantage associated with the use of filament composites is that their structural properties may be tailored to the expected loads in different directions. Contrary to metals which have the same material properties in all directions, filament composites are strongest in the direction the fibers are running. If a structural element such as a spar is to carry substantial load in only one direction, all the fibers can be oriented in that direction. This characteristic of filament composite provides for exceptional strength-to-weigh ratios and offers a tremendous weight savings opportunity to structural designers.
When fibers are aligned in only one direction, the resulting structure has maximum strength in that direction, and has little strength in other directions. Therefore, multiple layers or “plies” having fibers aligned in different directions with respect to one another are combined in a desired arrangement to provide combined strength along the principal axis as well as off-axis directions. As such, fibers oriented at 45° degree angles with the principle axis provide strength in two directions. For this reason, the 45° orientation is frequently used in structure that must resist torque. By utilizing permutations of this design philosophy to provide alternate plies of fibers at 0°, 45°, and 90° orientations the structural designer can obtain virtually any combination of tensile, compression, and shear strength in desired directions.
Common forms of fiber used in the production of composite structures include unidirectional tape, unidirectional fabric and bidirectional fabric. Unidirectional tape typically comes pre-impregnated with matrix material and is customarily provided on large rolls which can then be placed in a mold by hand or by robotic tape-laying machines. Similarly, bidirectional fabrics, having fibers running at 0 and 90 degrees, or unidirectional fabrics having fibers running in one direction may also be provided on large rolls pre-impregnated with matrix material. In another form of composite, individual filaments are wound around plugs or mandrels to form desired structural shapes. By way of background, the mandrels duplicate the inner skin of the structure or the inner surface of the structure. This technique is known as filament wound construction.
In addition to the form of fiber used in the production of composite structures, there are a number of fiber and matrix combinations which can be employed to provide desired structural properties of the resulting aircraft components. Fiberglass fiber embedded in an epoxy-resin matrix has been used for years for nonstructural components such as radomes and minor fairings. It is worthy of noting, however, that while fiberglass-epoxy has relatively good strength characteristics, its relatively low strength to weight ratio prevents its use in highly loaded structure. Additional material combinations which have eliminated this condition include: boron fibers used in combination with an epoxy matrix; aramid fibers (known as Kevlar) used in combination with an epoxy matrix, and graphite fibers used in combination with an epoxy matrix.
The United States military has been quick to incorporate fiber composite based structural components in its high-performance military aircraft. For example the F-16 utilizes graphite-epoxy composite material to form the horizontal and vertical tail skins. Similarly, graphite-epoxy composite material is utilized in the F/A-18 where such material forms the wing skins, the horizontal and vertical tail skins, the fuselage dorsal cover, the avionics bay door, the speed brake, and many of the control surfaces. The AV-8B employs composite materials even more extensively. In the AV-8B almost the entire wing, including the skin and substructure, is made of graphite-epoxy composite material with such material comprising approximately 26% of the total aircraft structural weight.
While composite materials have played an important role in reducing the overall structural weight of modem airframes, it should be noted that the basic design and layout of primary load carrying components contained within these structures has remained relatively the same. For example, a conventional aircraft wing structure consists of individual components such as spars, ribs, stringers and skin sections joined in combination to provide an integrated load carrying body which is capable of reacting to aerodynamic forces encountered during flight. As a result, individual spars, ribs, stringers and skin sections are specifically sized and oriented relative to one another so as to provide an optimized structural assembly designed to efficiently carry localized stresses generated by the combined effects of lift, drag, wind gusts, and acceleration loads which interact with surface of the wing or other airframe components.
In order to take advantage of weight savings opportunities afforded by the use of lighter weight materials, individual spars, ribs, stringers and skin sections previously formed from metal alloys have been replaced by similar components formed of fiber composite material. Frequently, these lighter weight components incorporate a “sandwich” style construction having two face sheets, or skins, made of fiber composite material which are bonded to and separated by a core. Typically, sandwich structures are formed with fiberglass-epoxy or graphite-epoxy skins which are bonded with adhesive to a phenolic honeycomb or rigid foam core wherein the skins carry tension and compression loads due to bending and the core carries shear loads as well as the compression loads perpendicular to the skins.
Unfortunately, manufacturing complexity and related labor cost associated with the assembly of numerous individual components, joined together to form an integrated load carrying structure, still remains. For example, conventional airframe construction techniques employ the use of elaborate jig fixtures designed to hold individual component parts in relative alignment during assembly to ensure proper component installation. In addition, drill templates are utilized to locate and drill fastener holes through mating pieces of structure to accommodate bolts or rivets used to mechanically join components together. These construction techniques are time consuming and require a great deal of dimensional precision because an improper installation of structural components may create a weakened resulting structure. Furthermore, the utilization of mechanical fasteners significantly contributes to overall structural weight. It is therefore generally desirable to minimize the number of mechanical joints in a structure in order to minimize both its weight and manufacturing cost while ensuring structural integrity. Integrally formed fiber composites structures have an important advantage over complicated structural assemblies in this respect, since large one-piece components are readily produced.
What has been needed and heretofore unavailable is a one-piece structure which is integrally formed as a unitary body and which is optimized to efficiently carry localized stresses developed from the complex interaction of static and aerodynamic forces encountered during all aspects of aircraft operation. The present invention satisfies these needs.